Turbine engines may include a compressor section, wherein inlet air is compressed, followed by a combustor section wherein fuel is combusted with the compressed air to generate exhaust gas. The exhaust gas is then directed to a turbine section, wherein energy is extracted from the exhaust gas.
The turbine section may comprise a rotor assembly. The rotor assembly may include a plurality of turbine blades installed on a rotatable disk. During operation, the turbine blades, the rotating disk, and other components of the turbine section may be exposed to elevated gas path temperatures, and thus may require cooling. Cooling may be provided to turbine section components using cooling air extracted from other parts of the engine via ducts in stationary and/or rotating components. For example, cooling air may be supplied from the combustor plenum. When the cooling air comes onboard the turbine rotor energy losses may occur resulting in a drop in pressure and a temperature increase. To minimize these effects, the static-to-rotating transition is typically accomplished using a stationary Tangential OnBoard Injector (TOBI).
A typical TOBI has a series of circumferentially spaced nozzle orifices which accelerate and direct the cooling air via a plurality of openings such that its tangential speed matches or exceeds that of the rotating components at the radius where the flow is being introduced. This speed matching approach typically reduces aerodynamic losses. A plurality of cooling holes serves as an inlet for cooling air in fluid communication with the turbine blade feed system. A portion of the cooling air leaving the TOBI openings traverses the distance between the TOBI exit and enters the plurality of cooling holes, which ultimately provides cooling to the turbine blades.
The cooling effectiveness of the blade is a strong function of the amount of cooling air, temperature of the cooling air, and the pressure level at which the cooling air is supplied to the blade. Increased flow rates of cooling airflow combined with elevated coolant supply pressure can be used to improve blade cooling effectiveness. The greater the cooling air flow, the cooler the turbine blade. However, gas turbine engines must be designed such that the use of secondary airflow is minimized, as that secondary cooling flow is diverted from the main core flow and does not do any actual work in the upstream turbine stages where it is dumped. Similarly, colder cooling air will more effectively cool the turbine blades. Accordingly, the preferable approach to cooling turbine blades is to use colder air, but less of it. The challenge lies in transferring the air from the static structure to the rotating components while minimizing the pressure drop and temperature increase.
In general, there are two types of TOBI configurations employed in gas turbine engines: radial and axial. In a radial TOBI, a cooling air flow comes onboard the turbine rotor at a low radius, and is then pumped to a higher radius where the blade slot bottom resides. A radial TOBI is the simpler and less expensive type of TOBI to manufacture, as well as the easiest to seal against leakage. Due to the inherent lower disk cavity pressures at low radii, the radial TOBI can achieve a higher pressure drop (i.e. higher swirl ratio) leading to a lower exit relative temperature. A primary weakness with the radial TOBI is that by pumping the cooling air to a higher radius where the blade slot bottom resides, as in a centrifugal compressor, the temperature of the air is raised significantly and the swirl ratio is lowered by the time it reaches the blade. A positive aspect of this type of TOBI configuration is that much of the pre-TOBI pressure is recovered.
In an axial TOBI configuration, the cooling air is introduced to the rotor at a higher radius to avoid heat gain and work extraction. In that there is no pressure recovery, the maximum TOBI swirl ratio is limited and leakages are higher. However, because there is no temperature increase, the overall turbine stage and cycle efficiency is better than in a radial TOBI.
Many small turbine engines, which spin in the 30,000 to 50,000+ RPM range, use axial TOBI setups at low radius where the cooling air passes through holes in a turbine rotor forward seal plate in order to reach the blade slot bottom at higher radius. This is necessary due to the high rotational speeds and high disk stresses. In this type of configuration, the rotor seal plates must be self-supporting and thus mandates that the engine have a low radius hub to carry the centrifugal load and have sufficient stress rupture life and burst margin. Larger, slower turning engines may employ a direct transfer TOBI at a higher radius where there is no need for a rotor seal plate. A direct transfer TOBI is the most favorable configuration in terms of minimizing secondary cooling flow and maximizing turbine efficiency, however this type of configuration is not feasible for small, high-performance turbines due to the inherent high rotor stresses precluding rotor holes feeding the blade slots.
Hence, there is a need for an apparatus including a direct transfer axial TOBI configuration for small turbine engines employing a rotor seal plate that provides for maximum temperature reduction with minimum pressure drop. In addition there is a need for a direct transfer axial TOBI configuration that does not adversely impact gas turbine engine efficiency, and/or does not adversely impact overall operational efficiency and cost. The present invention addresses one or more of these needs.